The present invention is directed to gas turbine engine architecture designs. In particular, the invention is directed to turbofan engines designed for supersonic and sub-sonic flight.
As gas turbine engines have evolved to achieve higher flight speeds, the temperatures and speeds within the high spool have proportionately risen. With current technology, it is becoming difficult or impossible to increase the pressure ratio within the high spool at such speeds and temperatures to further increase efficiency. In particular, the T3 temperature at the inlet of the combustor and the T4 temperature at the inlet of the high pressure turbine have risen to levels that produce unacceptably high stress levels and creep limitations in rotating components. There is a need for systems and methods that enable pressure ratios within the high pressure spool to be increased beyond current levels to further increase engine efficiencies at supersonic and sub-sonic speeds.